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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Effects of Forebody Deflection on the Stability and Control Characteristics of a Canard Airplane Configuration with a High Trapezoidal Wing at a Mach Number of 2.01

An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 to determine the effects of forebody deflection on the stability and control characteristics of a canard airplane configuration. The configuration had a high trapezoidal aspect-ratio-3 wing, a trapezoidal canard surface, and a single swept vertical tail. Forebody deflection angles of 0 deg, 2 deg and deg were investigated. The results indicated that nose-up deflections of the forebody provided positive increments of pitching moment with little increase in drag and hence would be useful in reducing the pitch-control requirements and the attendant losses in lift-drag ratio due to trimming. Deflection of the forebody, however, aggravated the decrease in directional stability with increasing angle of attack by causing a loss in tail contribution and by increasing the instability of the wing-body combination.

Spearman, M. Leroy↗
NASA NTRS2000

Numerical Study of Rarefied Hypersonic Flow Interacting with a Continuum Jet

An uncoupled Computational Fluid Dynamics-Direct Simulation Monte Carlo (CFD-DSMC) technique is developed and applied to provide solutions for continuum jets interacting with rarefied external flows. The technique is based on a correlation of the appropriate Bird breakdown parameter for a transitional-rarefied condition that defines a surface within which the continuum solution is unaffected by the external flow-jet interaction. The method is applied to two problems to assess and demonstrate its validity; one of a jet interaction in the transitional-rarefied flow regime and the other in the moderately rarefied regime. Results show that the appropriate Bird breakdown surface for uncoupling the continuum and non-continuum solutions is a function of a non-dimensional parameter relating the momentum flux and collisionality between the two interacting flows. The correlation is exploited for the simulation of a jet interaction modeled for an experimental condition in the transitional-rarefied flow regime and the validity of the correlation is demonstrated. The uncoupled technique is also applied to an aerobraking flight condition for the Mars Global Surveyor spacecraft with attitude control system jet interaction. Aerodynamic yawing moment coefficients for cases without and with jet interaction at various angles-of-attack were predicted, and results from the present method compare well with values published previously. The flow field and surface properties are analyzed in some detail to describe the mechanism by which the jet interaction affects the aerodynamics.

Glass, Christopher E.↗
NASA NTRS1934

The calculation of lateral stability with free controls

The discussion of the structural methods for obtaining lateral stability discloses the remarkable influence of the constant fuselage and wing proportions to the yawing moments. For the effectiveness of modifications in vertical tail surfaces and tail length, these quotas - little observed heretofore, in this connection - are decisive. This also applies to the amount of dihedral of the wing with regard to the roll stability of the complete wing already existing without angle of the dihedral.

Mathias, Gotthold↗
NASA NTRS1983

Uniaxial aerodynamic attitude control of artificial satellites

Within the context of a simple mechanical model the paper examines the movement of a satellite with respect to the center of masses under conditions of uniaxial aerodynamic attitude control. The equations of motion of the satellite take account of the gravitational and restorative aerodynamic moments. It is presumed that the aerodynamic moment is much larger than the gravitational, and the motion equations contain a large parameter. A two-parameter integrated surface of these equations is constructed in the form of formal series in terms of negative powers of the large parameter, describing the oscillations and rotations of the satellite about its lengthwise axis, approximately oriented along the orbital tangent. It is proposed to treat such movements as nominal undisturbed motions of the satellite under conditions of aerodynamic attitude control. A numerical investigation is made for the above integrated surface.

Sazonov, V. V.↗
DOE OSTI2022

Thermophysical properties of FLiBe using moment tensor potentials

Fluoride salts are prospective materials for applications in some next-generation nuclear reactors and their thermophysical properties at various conditions are of interest. Experimental measurement of the properties of these salts is often difficult and, in some cases, unfeasible due to challenges from high temperatures, impurity control, and corrosivity. Therefore, accurate theoretical methods are needed for fluoride salt property prediction. In this work, we used moment tensor potentials (MTP) to approximate the potential energy surface of eutectic FLiBe (66.6% LiF – 33.3% BeF2) predicted by the ab initio (DFT-D3) method. Here, we then used the developed potential and molecular dynamics to obtain several thermophysical properties of FLiBe, including radial distribution functions, density, self-diffusion coefficients, thermal expansion, specific heat capacity, bulk modulus, viscosity, and thermal conductivity. Our results show that the MTP potential approximates the potential energy surface accurately and the overall approach yields very good agreement with experimental values. The converged fitting can be obtained with less than 600 configurations generated from DFT calculations, which data can be generated in just 1200 core hours on today's typical processors. The MTP potential is faster than many machine learning potentials and about one order of magnitude slower than widely used empirical molten salt potentials such as Tosi/Fumi.

36 MATERIALS SCIENCE↗
NASA NTRS1945

Summary of Airfoil Data

Recent airfoil data for both flight and wind-tunnel tests have been collected and correlated insofar as possible. The flight data consist largely of drag measurements made by the wake-survey method. Most of the data on airfoil section characteristics were obtained in the Langley two-dimensional low-turbulence pressure tunnel. Detail data necessary for the application of NACA 6-serles airfoils to wing design are presented in supplementary figures, together with recent data for the NACA 24-, 44-, and 230-series airfoils. The general methods used to derive the basic thickness forms for NACA 6- and 7-series airfoils and their corresponding pressure distributions are presented. Data and methods are given for rapidly obtaining the approximate pressure distributions for NACA four-digit, five-digit, 6-, and 7-series airfoils. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed. The data indicate that the effects of surface condition on the lift and drag characteristics are at least as large as the effects of the airfoil shape and must be considered in airfoil selection and the prediction of wing characteristics. Airfoils permitting extensive laminar flow, such as the NACA 6-series airfoils, have much lower drag coefficients at high speed and cruising lift coefficients than earlier types-of airfoils if, and only if, the wing surfaces are sufficiently smooth and fair. The NACA 6-series airfoils also have favorable critical-speed characteristics and do not appear to present unusual problems associated with the application of high-lift and lateral-control devices. Much of the data given in the NACA Advance Confidential Report entitled "Preliminary Low-Drag-Airfoil and Flap Data from Tests at Large Reynolds Number and Low Turbulence," by Eastman N. Jacobs, Ira R. Abbott, and Milton Davidson, March 1942 has been corrected and included in the present paper, which supersedes the previously published paper.

LOW-DRAG↗
NASA NTRSDate not supplied

Towards Plume Impingement Modeling in Space Environments

After 30 years human presence in low-earth orbit, NASA is returning to the moon and eventually will go to Mars. To this end, NASA is constructing the Lunar Gateway, an ISS-like space station to act as a home base for Lunar exploration. A large space station must be assembled while in orbit, using a “piecemeal” approach. In this context, it means that the different modules will arrive at different times and attach to what is already in service. The ISS provides an excellent example of this approach, and the proposed Lunar Gateway will undergo a similar assembly process. This assembly is achieved via “docking” maneuvers between modules, which are made possible by sequential firings of the onboard reaction control system (RCS) thrusters. They work by firing hot gases to produce adverse thrust and the needed change in velocity to safely finish the docking approach. The issue is that the gas from these thrusters' forms flow structures described as “plumes” and can impinge onto the outer surfaces of the space station, causing unwanted forces and moments, heat loads, sediment deposition, and in extreme cases, even surface erosion. All mechanisms that can damage the space station and must be avoided. Both permanent and visiting modules will have these RCS thruster exhaust impingement problems. Accurately and efficiently modeling these plume is an involved multi-physics calculation but also an important tool when designing the control algorithms of approaching modules. This poster presents progress towards this simulation on two fronts. First is the verification of OpenFOAM for rarefied plume impingement calculations by direct comparison to published DAC cases. Second is the estimation of plume impingement strikes over the time scale of an entire docking event. This is done by using a simple plume source flow model and a prescribed motion visualizer—coded in Python. Together these tools help push NASA's capabilities for simulating these plume impingement effects.

Rarefied Flows↗
NASA NTRS1975

Mated aerodynamic characteristics investigation for 0.04-scale model Boeing 747 CAM/external tank (model AX1284 E-5) combination in the University of Washington Aeronautical Laboratory F. K. Kirsten Wind Tunnel (CA11)

Experimental investigations of the aerodynamic characteristics of a 0.04-scale external tank (ET) force model in combination with a 0.04-scale Boeing 747 force model were conducted. Test purposes were: (1) to determine ET airloads for selected configurations and (2) to determine the effectiveness of ET position, incidence, and support structure and 747 vertical stabilizing surfaces on stability, control, and performance of 747/ET combinations. The 747 was tested alone to establish baseline data and to verify test results. Six-component aerodynamic force and moment data were recorded for the 747 CAM and ET combination. Six-component force and moment data were also recorded for the ET, which was mounted on an internal balance supported by the 747. Data were recorded for angles of attack from -4 deg to +24 deg in 2 deg increments and angles of sideslip of - deg to + or - 20 deg. Testing was conducted at Mach 0.15 with dynamic pressure deg at 36 psf and unit Reynolds number of 1.3 million per foot. Photographs of test configurations are shown.

Source record↗
NASA NTRS1987

Exploratory studies of actuated forebody strakes for yaw control at high angles of attack

Wind-tunnel studies have been conducted to evaluate the potential of actuated forebody strakes to provide increased levels of yaw control on fighter aircraft at extremely high angles of attack where conventional aerodynamic controls are ineffective. The studies involved low-speed wind-tunnel tests of actuated forebody strake concepts applied to a generic fighter model and included circumferential pressure and flow visualization surveys on an isolated forebody. Results showed that the actuated forebody strake concept can provide high levels of yaw control over wide ranges of angle-of-attack and sideslip. However, when lifting surfaces were placed in close proximity to the forebody/strake combination, significant interaction effects occurred which reduced the yaw control effectiveness of the strakes and induced coupled rolling and pitching moments.

Murri, Daniel G.↗
NASA NTRS1993

Subsonic aerodynamic characteristics of a proposed advanced manned launch system orbiter configuration

The Advanced Manned Launch System is a proposed near-term technology, two-stage, fully reusable launch system that consists of an unmanned glide-back booster and a manned orbiter. An orbiter model that featured a large fuselage and an aft delta wing with tip fins was tested in the Langley 7- by 10-Foot High-Speed Tunnel. A crew cabin, large payload fairing, and crew access tunnel were mounted on the upper body. The results of the investigation indicated that the configuration was longitudinally stable to an angle of attack of about 6 deg about a center-of-gravity position of 0.7 body length. The model had an untrimmed lift-drag ratio of 6.6, but could not be trimmed at positive lift. The orbiter model was also directionally unstable. The payload fairing was responsible for about half the instability. The tip-fin controllers, which are designed as active controls to produce artificial directional stability, were effective in producing yawing moment, but sizable adverse rolling moment occurred at angles of attack above 6 deg. Differential deflection of the elevon surfaces was effective in producing rolling moment with only small values of adverse yawing moment.

Ware, George M.↗
NASA NTRS2012

Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA)

With the very recent advent of NASA's Environmentally Responsible Aviation Project (ERA), which is dedicated to designing aircraft that will reduce the impact of aviation on the environment, there is a need for research and development of methodologies to minimize fuel burn, emissions, and reduce community noise produced by regional airliners. ERA tackles airframe technology, propulsion technology, and vehicle systems integration to meet performance objectives in the time frame for the aircraft to be at a Technology Readiness Level (TRL) of 4-6 by the year of 2020 (deemed N+2). The proceeding project that investigated similar goals to ERA was NASA's Subsonic Fixed Wing (SFW). SFW focused on conducting research to improve prediction methods and technologies that will produce lower noise, lower emissions, and higher performing subsonic aircraft for the Next Generation Air Transportation System. The work provided in this investigation was a NASA Research Announcement (NRA) contract #NNL07AA55C funded by Subsonic Fixed Wing. The project started in 2007 with a specific goal of conducting a large-scale wind tunnel test along with the development of new and improved predictive codes for the advanced powered-lift concepts. Many of the predictive codes were incorporated to refine the wind tunnel model outer mold line design. The large scale wind tunnel test goal was to investigate powered lift technologies and provide an experimental database to validate current and future modeling techniques. Powered-lift concepts investigated were Circulation Control (CC) wing in conjunction with over-the-wing mounted engines to entrain the exhaust to further increase the lift generated by CC technologies alone. The NRA was a five-year effort; during the first year the objective was to select and refine CESTOL concepts and then to complete a preliminary design of a large-scale wind tunnel model for the large scale test. During the second, third, and fourth years the large-scale wind tunnel model design would be completed, manufactured, and calibrated. During the fifth year the large scale wind tunnel test was conducted. This technical memo will describe all phases of the Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA) project and provide a brief summary of the background and modeling efforts involved in the NRA. The conceptual designs considered for this project and the decision process for the selected configuration adapted for a wind tunnel model will be briefly discussed. The internal configuration of AMELIA, and the internal measurements chosen in order to satisfy the requirements of obtaining a database of experimental data to be used for future computational model validations. The external experimental techniques that were employed during the test, along with the large-scale wind tunnel test facility are covered in great detail. Experimental measurements in the database include forces and moments, and surface pressure distributions, local skin friction measurements, boundary and shear layer velocity profiles, far-field acoustic data and noise signatures from turbofan propulsion simulators. Results and discussion of the circulation control performance, over-the-wing mounted engines, and the combined performance are also discussed in great detail.

Subsonic Fixed Wing↗
DOE OSTI2023

Utah FORGE: Fault Shear Reactivation Experimental Data for Fluid Injection-Rate Controls on Seismic Moment

Included are experimental data recorded from shear experiments that specifically explore the link between fluid-injection rate and seismic moment resulting from shear reactivation of laboratory faults. Raw mechanical data from three experiments are included alongside corresponding MATLAB scripts that import and plot the data, as well as use it to calculate shear and normal stress. Experiments are performed on 2.5-3 inch long granitoid cores from the Utah FORGE EGS demonstration site, containing a single inclined fracture with small-scale roughness added to the fracture surface. The raw data included here were recorded from an aluminum triaxial pressure vessel (TEMCO) configured with three independent servo-controlled pumps, with distilled water used as the working fluid. The pumps control confining pressure, upstream pore pressure, and axial pressure, with each pump connected to a LabView interface to record applied pressures, cumulative injected water volumes, and pump flow rates. The downstream outlet from the fracture is closed to allow pressurization, which is measured by an external pressure transducer. A linear variable differential transformer (LVDT) attached to the axial piston measures axial displacement, from which we calculate shear displacement along the fracture. Additionally, P-wave transducers are used to record acoustic signatures, where acoustic emission events and maximum amplitudes are compared against seismic moment and shear slip velocity. Fluid injection rates range between 0.05 mL/min, 0.25 mL/min, and 0.75 mL/min for each experiment. Along-fault pressure distributions are progressively less uniform as injection rates increase, representing a switch from steady-state to transient conditions. Triggered shear displacement is used as a proxy for seismic moment and is indexed against cumulative injection volume and rate. Each experiment is performed under constant shear stress conditions, and the sample is fully saturated with DI water. Axial and confining stresses are applied to 3 MPa through pressure-stepping in 500 kPa increments. The pore pressure is held constant at 200 kPa prior to initiating the experiment, and initial axial displacement is recorded. The axial stress is then increased to initiate shear mobilization during the loading phase (run-in) until a peak steady state is achieved. The initial shear stress is reduced to approximately 80% of the peak shear stress by decreasing the axial stress, then held constant for the duration of each experiment.

15 GEOTHERMAL ENERGY↗
NASA NTRS1978

Estimate of Orbiter static aeroelasticity properties via FLEXSTAB

The theory underlying the use of rigid-model experimental data in FLEXSTAB aeroelasticity calculations is summarized. A description is given of the aerodynamic and structural models used in the calculation of Space Shuttle Orbiter aeroelasticity properties during the re-entry phase. Experience gained with the use of pressure, nonlinear force and moment, and area-ratio correction methods in the Orbiter application is described and typical program output data are included. The inability of the program to correctly estimate differential control surface pressures and to accept external corrections for these terms is shown to be a major limitation in the Orbiter application.

Cavin, R. K., III↗
NASA NTRS2017

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Erickson, Gary E.↗
NASA NTRS1989

Nonprincipal plane scattering of flat plates and pattern control of horn antennas

Using the geometrical theory of diffraction, the traditional method of high frequency scattering analysis, the prediction of the radar cross section of a perfectly conducting, flat, rectangular plate is limited to principal planes. Part A of this report predicts the radar cross section in nonprincipal planes using the method of equivalent currents. This technique is based on an asymptotic end-point reduction of the surface radiation integrals for an infinite wedge and enables nonprincipal plane prediction. The predicted radar cross sections for both horizontal and vertical polarizations are compared to moment method results and experimental data from Arizona State University's anechoic chamber. In part B, a variational calculus approach to the pattern control of the horn antenna is outlined. The approach starts with the optimization of the aperture field distribution so that the control of the radiation pattern in a range of directions can be realized. A control functional is thus formulated. Next, a spectral analysis method is introduced to solve for the eigenfunctions from the extremal condition of the formulated functional. Solutions to the optimized aperture field distribution are then obtained.

Balanis, Constantine A.↗
NASA NTRS1991

Unsteady surface pressure measurements on a slender delta wing undergoing limit cycle wing rock

An experimental investigation of slender wing limit cycle motion known as wing rock was investigated using two unique experimental systems. Dynamic roll moment measurements and visualization data on the leading edge vortices were obtained using a free to roll apparatus that incorporates an airbearing spindle. In addition, both static and unsteady surface pressure data was measured on the top and bottom surfaces of the model. To obtain the unsteady surface pressure data a new computer controller drive system was developed to accurately reproduce the free to roll time history motions. The data from these experiments include, roll angle time histories, vortex trajectory data on the position of the vortices relative to the model's surface, and surface pressure measurements as a function of roll angle when the model is stationary or undergoing a wing rock motion. The roll time history data was numerically differentiated to determine the dynamic roll moment coefficient. An analysis of these data revealed that the primary mechanism for the limit cycle behavior was a time lag in the position of the vortices normal to the wing surface.

Arena, Andrew S., Jr.↗
NASA NTRS2001

Recent CFD Simulations of Shuttle Orbiter Contingency Abort Aerodynamics

Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20-60 degrees, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). While approximately 40 cases have been computed, only a sampling of the results is presented here. The computed results, in general, are in good agreement with the Orbiter Operational Aerodynamic Data Book (OADB) data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects.

Papadopoulos, Periklis↗
NASA NTRS1980

Aerodynamic and inlet flow characteristics of several hypersonic airbreathing missile concepts

Four conceptual hypersonic missile configurations were examined experimentally and theoretically. Two of the concepts employed twin module bottom-mounted engines and two were designed for upper surface inlets or engines with the intent of reducing the vehicle observables. The tests were conducted at Mach 6 and Reynolds numbers of 6 to 7.5 x 10 to the 6th per foot. Flow field surveys in the vicinity of the engine inlet were made on all configurations and force and moment tests were conducted on three of the vehicles. Stability and control characteristics of the bottom-mounted engine configurations which incorporated slender, low wings were dominated by strong vortices that promoted severe pitchup tendencies. The shock layer and flow quality in the vicinity of the bottom-mounted engine inlets were dependent on nose shape. The spatula-like upper surface engine concept demonstrated good performance and had uniform flow entering the engine inlet, while the upper surface inlet concept with a highly swept forebody incurred large gradients due to interactions with leading edge shocks.

Dillon, J. L.↗